Aircraft take-off systems



Jan. 3, 1967 E. PRIES'TLEY AIRCRAFT TAKE-OFF SYSTEMS 5 Sheets-Sheet lFiled April 17, 1964 /NvE/vrop E c. PRIESTLE R BY y t t M ATTORNEYS Jan.3, 1967 E, PRIESTLEY AIRCRAFT TAKE-OFF SYSTEMS 5 Sheets-Sheet 2 FiledApril 17. 1964 Jan. 3, 1967 E. PRlr-:sTLEY 3,295,359

AIRCRAFT TAKE-OFF SYSTEMS Filed April 17, 1964 s sneetssheet a v/NVENToR ERC PRESTLY 3y n YM ATTORNEYS United States Patent O AIRCRAFTTAKE-OFF SYSTEMS Eric Priestley, Elliott Brothers (London) Limited, St.Leonards, Hastings, Sussex, England Filed Apr. 17, 1964, Ser. No.360,536 Claims priority, application Great Britain, Apr. 19, 1963,15,636/ 63 16 Claims. (Cl. 73-178) This invention relates to an aircrafttake-off director system for affording to the pilot of an aircraft acontinuous indication of the optimal action to be taken during theperiod of take-off and initial climb. The director system may also bearranged to control an automatic pilot.

The task of an aircraft pilot during take-off and initial climb is tostart nose-up rotation of the aircraft at an indicated airspeed VR,rotate as rapidly as reasonably possible, without scraping the tail-skidon the runway and, after lift-01T, i.e. becoming airborne, to pull theaircraft up quickly to attain at least screen height (35 feet) at theend of the available take-off distance, at a speed not less than thetake-ott safety speed, and into the take-off climb, at speed VTOC, in astable manner.

Possible complicating factors which may arise are as follows- (a) t'hesmall amount of lift available for pull-up (b) necessity to obtainadequate -clearance over obstacles (c) necessity to turn to avoid highground (d) existence of a degree of spiral instability throughout thetake-off period (e) reduction of lap angle at or above 400 feet (f)occurrence of subnormal thrust, or of thrust reduction, at any ltimeduring take-off and initial climb (g) implementation of noise abatementprocedure.

The pilots ltask following a decision to abort a landing is to make asafe pull-up into what is normally referred to as approach climb.

Possible complicating factors in this case are- (i) time lag betweenopening throttles and obtaining the corresponding extra thrust (ii)necessity to reduce flap angle from approach or landing setting totake-off setting;

together with some or all of (a) to' (g) above.

To assist the pilot in these complex manoeuvres an important object ofthe invention is to provide an aircraft take-off director system inwhich certain components of the aircrafts movement and position aremeasured, signals corresponding to the measurements are produced,further signals are generated representing constants which arecharacteristic of the aircraft, and the measurements signals and theconstants signals are combined order .to give to the pilot continuousinformation as to the action to be taken at the controls in order toprovide the optimal conditions of performance and safety throughout theperiod of take-off and initial climb.

Another object is to provide a director system in which the pilot isgiven a clear indication on a single indicating director instrument ofthe action to be taken throughout the period of take-off and initialclimb.

A further object is to provide a director system which provides thepilot with all the information he needs on the behavior of the aircraftduring the take-off and initial climb period.

A still further object is to provide a new director system for use bythe pilot during take-off and initial climb.

Still another object is to provide a director system which may be usedby the aircraft pilot and which may also be used to control an automaticpilot.

An additional object is to provide a director system 3,295,369 PatentedJan. 3, 1967 by means of which the take-off and initial climb proceduremay be carried out in accordance with a predetermined programme based onthe characteristics of the aircraft.

Broadly, a take-off director system for aircraft according to :theinvention comprises means for continuously measuring selected componentsof the movement and position of the aircraft, means to generate signalsrepresenting constants dependent upon .the characteristics of theaircraft, means to combine measurements signals with the constantssignals to provide a composite demand signal representative of theaction required to perform the take-off and initial climb cycle inaccordance with a predetermined programme, and a director indicatorinstrument .to display the composite signal so that the pilot, bysetting the controls to carry out `the demands shown by the system, canperform lthe required sequence of takeoff control operations in asubstantially optimal manner. The director display instrument may bearranged to provide a two-dimensional indication, so that spiralinstability is shown and may be corrected. AIn case of a reduction inthrust the director indicating instrument will continue to display theoptimal demand appropriate to the special circumstances.

The invention includes means associated with the main undercarriage at aselected .point in the take-off procedure such that when the aircraftbecomes airborne the fact is signalled by the undercarriage oleo legs.becoming fully extended, and the signal is fed into the system, so thatthe system takes into account the fact that the aircraft has becomeairborne.

`In addition to providing 'the human pilot with demand information, thedirector system of the invention may be used to control an automaticpilot.

Other objects will appear from the following description of twoembodiments of the director system according to the invention, given byway of example, with reference to the accompanying drawings, in which-FIGURE 1 (not to scale) shows -diagrammatically the take-off and initialclimb of an aircraft, broken down into fou-r sequential phases;

FIGURE 2 is a block schematic diagram of a simple basic form of take-offdirector system according to the invention;

FIGURE 3 is a block schematic diagram of a more complex form of take-offdirector system according to the invention; and

FIGURE 4 is a view of the face of the director indicating instrument.

Referring to the drawings, FIGURE 1 is a diagram showing the take-offand initial climb of an aircraft, carried out in accordance ywith -apredetermined programme, broken down into four sequential phases, lasfollows:

Phase A (pre-rotation) extends from the start of the take-off ground runof the aircraft 11, up to the point indicated by reference 12, with theaircraft: in the position 11a, at which the aircraft has attained anairspeed VR, scheduled for the beginning of rotation.

Phase B (rotation) consists of the progressive nose-up rotation of theaircraft with the main wheels still on the runway, until lift-off occurswhen the aircraft has reached the position 13 on the runway, and itsattitude is as ndicated at 11b.

Phase C (pull-up) after passing the point 13, the aircraft lifts off andits incidence is further increased to produce upward accelerationthrough the attitude indicated at 11C, at which the undercarriage isalmost fully retracted, the acceleration being progressive reduced asthe aircraft approaches the scheduled take-olf climb speed VTOC. At thispoint the aircraft has reached the point 14 and its attitude is asindicated-by 11d.

Phase D (initial climb) is the substantially constantspeed climb up to aheight of the order of 1500 feet, at

point 15, the attitude of the aircraft being as indicated at 11e, whichis the same as that shown at 11d.

In the system shown in FIGURE 2 the demands signalled to the pilotrequire the use of only two measuring instruments. These are anintegrating pitch rate gyroscope 16 having one output q and a secondintegrated output fq, and an airspeed sensor 17 having an output V.

FIGURE 2 also shows the other units included in the system. The otherunits are a time delay unit 1&8 which provides a delay T1, which ispreset; a limit constant unit 109 which provides a limit constant k1,which is preset; a take-off climb-speed unit 2h which provides a signalVTOC, which is preset; a rotation speed unit 18 to provide a signal VR,which is preset; a rotation warning speed unit 19 to provide a signalVRW, which is preset; and a pitch-up demand unit 55 to provide a signalAHR, which is present.

FIGURE 2 also shows a voltage source 23 feeding switches, respectively24 and 24a, which are open until the oleo legs become fully extended asthe aircraft lifts olf the ground. The switches are wired in series sothat no signal is given until both legs are fully extended.

A phase advance unit 25 provides a time constant T2, and other unitscomprise a limiter 26, a shaping unit 27 and a Director on switch 28.

Further units c-ontained in the system are amplifiers, respectively 34,35, 36, 37 and 3S, and the system is completed by a three-way switch 4t)for the director indicator instrument 41.

The constants previously referred to which are employed in the systemaccording to FIGURE 2 are as follows.

T1, provided by the time delay unit 1498, provides a time delay, whichmay be of the order of two `seconds from the commencement of phase Bduring which the closure of switches 24 and 24a will not initiate phaseC. Phase B must therefore be at least T1 seconds in duration. q

The limit constant k1 provided by the unit 109 is a limiting constantwhich is applied to the limiter 26 to prevent the signal from the phaseadvance unit 25 from exceeding a certain amplitude. The reason is thatthis signal acts to demand a proportional nose-up pitch rate which isrequired not to exceed a certain constant value during the early or anypart of phase C (pull-up) lest the aircraft incidence should approachtoo close to the stall condition.

The take-olf climb-speed signal VTOC provided by the unit 20 correspondsto the correct intial clirnb-speed.

The variable signal V provided by the airspeed measuring device 17represents the instantaneous airspeed of the aircraft.

'Iihe constant VR provided by the unit 18 is the rotation speed which isscheduled for the particular aircraft.

The rotation warning speed signal VRW, provided by the unit 19 is toindicate that the aircraft speed is closely approaching rotation speed.

The pitch-up demand signal AGR provided by the unit S is chos-en to besomewhat greater than the nose-up change of attit-ude which will givelift-off.

The integrating pitch rate gyroscope needs no particular description,except to mention that integrating circuitry is provided so that a ratesignal q and a second signal fq are available.

The shaping unit 27 is provided for the purpose of holding the directorindex so that it cannot move more than a prescribed amount below theindicator zero pitch position, except when the aire-raft pitches nose-upat a speed less than VRW.

The phase -advance unit 25 is provided to give anticipation to theindicator index of the imminence of reaching VTOC, thereby indicatingthe need to reduce acceleration in good time. It also provides dampingthroughout the later stages of the take-olf.

FIGURE 4 shows the face of the director indicating instrument. Theinstrument itself vmay conveniently be constituted by a cathode ray tubehaving `a calibrated scale mounted in front of the screen face. Othertypes of indicator may, of course, be used. As shown in FIGURE 4, thescale is of square shape. At the lefthand side of the instrument is ascale of pitch rate demand, the figures above the zero line indicating anose-up demand scale-d in `degrees per second and the figures below thezero line indicating a nose-down or speed increase demand scaled inknots. The response to a nose-down demand under particular conditions,i.e., by the pilot setting his controls appropriately to put the nose ofthe aircraft down, will lead to an increase in speed so that a nose-downdemand effectively connotes a speed increase demand. 'I'.he -right-handportion Iof the instrument face above the zero line is calibrated inpitch angle change demand in -degrees and reads from zero to 15. Thescale contains la zero marker 56, and above it is a marker 57 whichindicates a pull-up pitch rate demand of 11/2 /sec., which is apredetermined limit. Below the Zero line 56 is a marker 58 whichindicates an airspeed VRW which, in the example being described, is l0knots less than VR.

In operation, the cathode ray spot is made to oscill-ate rapidly acrossthe instrument dial to produce a horizontal line. Where the systemincludes means to indicate spiral rotation of the aircraft the line willtilt to left or right as the aircraft banks to one side or the other, asillustrated, 4for example, by the line 60, which represents the aircraftbanked about 12 to starboard.

The sequence and method of operation of the system are as follows:

(l) Power on..-When power is switched on, the threeway switch 40 movesto `a first position in which the moving member 42 makes contact with afirst contact stud 43. The integrating pitch rate gyroscope 15 is run upto speed. The signal q from the gyroscope 16 is applied over a line 44to the amplifier 38, whose gain is kq. The output signal from theamplifier is -kqq and passes over a line 45 to the -director indicatinginstrument 41. The index of the indicator is immediately centered underthe action of the electrical spring of the gyroscope gimbal/ rotorsystem. When the rotor has run up to full speed the index moves aboutthe zero demand marker line with aircraft pitch rate.

(2) Direczor @n.Before or during taxying to the takeoff starting pointthe director system is switched on by closing switch 28. The airspeedsensor signal V from the unit 17 and the rotation speed signal VR fromthe unit 18, which are combined in a ditferencing unit 48, produce aslgnal V- VR which is passed, together with the rotation warning speedsignal VRW from the unit 19, to the shaping unit 27, producing a signalrepresenting a function f(V-VR), which is fed to the amplifier 36. Thefunction f(V-VR) is equal to V-VR except for V VRW when the value islimited to VRW-VR. The amplifier 36 has a gain kv so that a signal kV f(V- VR) is passed through the switch 28 and a line 49 to contact 43,through contact arm 42 and over a line Sil to the indicating instrument41.

The total signal to the indicator 41 is now The airspeed is bound to beless, at this stage, than the rotation warning speed VRW so that unlessq is smaller than the director index moves downwardly to the marker 58which, with the scaling used for index displacement per unit value ofspeed difference V- VR or V-VTQC, is displaced below centre by an amountcorresponding to the preset value of VR-VRW, typically about 10 knots.The index moves about this lower marker with aircraft pitching.

Phase A.-As the aircraft accelerates in its take-olf run, the directorindex still oscillates about this line due to pitching until theairspeed reaches VRW, after which the index moves upwardly towards thecentre of the indicator face at a rate proportional to the rate ofincrease of airspeed and reaches the centre point when V= VR at the endof the phase. Throughout this phase the demand is plainly nose-down,with a Warning of the imminence of reaching VR as soon as the airspeedexceeds VRW. With VRW set at l knots below VR, and assuming the meanacceleration between these speeds to be three knots per second, thewarning time is just over three seconds.

As in Director on the signal to the director is the three-way switch 40being still in position l.

Phase B A predetermined increase of incidence is now required and thisimplies an equal increase of pitch attitude when the aircraft is rollingalong a substantially ilat runway. Also, since the duration of rotationis only about 3 to 5 seconds, it is quite legitimate to ignore gyroscopedrift which, for a typical low grade instrument in this integratingmode, would be of the order of 1/10/ second (6/minute) or less.

As the airspeed reaches VR, -at point 12 in FIGURE 1, a change of signof the V- VR signal into the three-way switch 40 causes the switch tomove to position 2, in which the switch arm 42 engages a xed contact 51.

A AGR signal from the -unit 5S is combined with the integrated signal fqfrom the gyroscope 16 in a differencing unit 52 and fed to the amplifier37, which has a gain of k. The combined signal k6(A0R-fq) is fed over aline 53 to the contact 51 and thence to the switch arm 42 and over lineS0 to the indicator 41. This causes, in conjunction with the signal-cqq, the index to be displaced from centre by an amount proportional toso that it flicks upwardly to demand nose-up rotation of the rightamount MR at an initial rate kqAR Assuming that the pilot answers thisdemand by pulling the aircraft nose up to centre the index, the demandshown by the index falls off, and by the time the index is again steadyat the central position with the mean q at Zero the aircraft will haverotated stably upwards through the angle ASR.

When V=VR there is an additional operation, not indicated in FIGURE 2,whereby the integrating pitch rate gyroscope 16 is switched from therate mode (integrating capacitor charged with voltage representing ABRbut not in spring coil circuit) to the rate-i-I rate mode (capacitance`in series with spring coil).

The maximum director index displacement may be made to correspond toabout of pitch angle demand. With k=3ka this maximum displacement willalso correspond to 5/second pitch rate demand. So, taking a typical MRof 12, the index will move up to fourfths of the maximum at the start ofphase B and this is also shown as a demand for 4/second pitch rate.During this phase the aircraft continues to accelerate. If AGR has beencorrectly chosen the aircraft will leave the runway decisively but withvery little vertical acceleration.

Phase C.-After lift-off the aircraft incidence is further increased toproduce upward acceleration, in spite of the subsequent loss of lift dueto reduction of ground effect, the g being later reduced as the speedapproaches VTOC to give a smooth transistion to initial climb. Thesystem is designed to demand no more than a constant chosen pitch rate,which is essential since the -airspeed may be no more than about 15% to20% above the stall speed, and some of this margin must be regarded asan assurance in respect of drop of head component of wind.

Furthermore, this constant demand is held for most of the pull-up,possibly for at least 5 seconds, so that the total pull-up time isminimised.

In the example being described the phase is initiated by full extensionof both port and starboard main oleo legs, which close microswitches 24and 24a when fully extended. These microswitches are wired in series.Any temporary simultaneous oleo extensions, due to unevenness intherunway, for example, are ignored and initiation is prohibited until atleast T1 seconds after the commencement of the phase B, the time delayT1 being provided by the signal from the unit 103. The threeway switch40 is now moved from position 2 to position 3, in which the movingcontact arm 42 engages a third fixed Contact 54. Consequently, thek,(A0R-fq) part of the phase B signal is replaced by a new signal. Thisis made yup of the signal VTOC, derived from the unit 2t), combined withthe airspeed sensor signal V from the unit 17 in a dilferencer unit 47to provide a signal V-T/T. This is amplied in amplifier 34,phaseadvanced in unit 25 to produce a resultant signal kv(Tz/+V-V'roc)and fed to the limiter 26, which also receives the signal kl from thelimit constant unit 109. The output signal from the limiter 26 is[kV(T2V'+V-VTOCI)] k1. Thus the pull-up is demanded by displacing thedirector index upwards by an amount proportional to k (Ta'f-x-V-VTOC) Qshould be 0f the order l v zvwarv-VTOC) te. -:ia TF6 l This demand willbe limited to some value of the order 11/2/second which in a steadypull-up (incidence constant) at a true airspeed of l2() knots givesabout 0.16 g.

With V of the order 2 knots/second or more the 11/2/ second limitershould operate initially, notwithstanding that usually V VTOC at thistime, the effect being to give an upward director index displacement tothe pullup demand limit, which will be pulled lby `the pilotestablishing a pitchf rate of l1/2/second until the phase advanced speeddemand signal [kV(T2V+V-Vfroc)l k1 comes ot its limit. When this happensthe pilot is commanded to take off the pitch rate and this should andnulling of this by centering the index will give stable take-off climbat airspeed VTOC.

It -Will be appreciated that the phase-advanced speed mode can readilybe stabilised by displacing the index laterally by lan amountproportional to bank angle and requiring the pilot to centre it.

This phase extends to perhaps 1,500 feet.

It will be appreciated that the phase-advanced speed error term TZV-|-V-VTOC, used in this and the previous phase may, in fact, lbe of morecomplex form.

Ejfect of thrust Iss.*lf thrust `loss occurs before liftoff nodiscontinuities in demand occur but the pull-up will automatically be ofshorter duration 'and the initial climb will `be at a atter angle. Ifthe loss occurs very soon after lift-off there may again be no obviouschange in the time pattern of demand. Later losses will result in adownward movement of the index in sympathy with the reduction in V, andin the take-off climb the resulting pilots action of pushing the .stick"forward to .centre the index will check the reduction in airspeed whichmight otherwise occur. It will be appreciated that provision for varyingVTOC if a thrust loss occurs may readily ybe made iby means of thrustsensors and logical units.

FIGURE 3 shows a more complex system which takes into account additionalcomponents of the aircrafts position and attitude and certain additionalconstants, all of which are desirable in a system intended for a largeaircraft. In addition to the pitch rate gyroscope 16 providing the q andf q signals and the airspeed sensor 17 to provide the signal V, there isa barometric height sensor 61 having an output h andan attitudereference system 62 (which also includes gyroscopes) having outputs 1]/(azimuth or heading angle) 0 (pitch angle) and gb (bank angle). Theattitude reference system is normally fitted in the aircraft in anycase, and is made use of in the system according to the invention.

A unit 63 is provided by which the pilot can set in the actual all-upweight of the aircraft at the beginning of the trip. This provides asignal W. A further unit is a ap angle sensor unit 64 which provides asignal corresponding to the actual setting of the aps. In the simplersystem of FIGURE 2, previously described, the flap angle was assumed tobe constant. In the present system, however, account is taken of apangle variation. This signal is combined with the signal representingthe all-up weight of the aircraft in a further unit 65 which maycomprise a twoand/or three-dimensional cam arrangement or its electronicequivalent. The unit 65 provides the signal VR, which is fed to adifferencing unit 66 where it is combined with the signal V from theairspeed sensor 17 to provide the signal V-VR which is fed to a limiter67 and will `be referred to more particularly later. The unit 65 alsoprovides signals over lines 68 to the unit 63 which includes means todisplay signals corresponding to VR, V2 and V3, the latter two of whichwill be explained later.

A unit 69 contains the pilots noise-abate switch for the purpose ofreducing the engine speed to keep the aircrafts noise within acceptablelimits. When this switch is closed the unit 69 changes the values of aconstant signal k1 supplied by a unit 70, and a further constant signalk2 supplied by a unit 71.

A unit 72 provides a constant k which is equal to VR-VRW, the constant[c4 being also applied to the limiter 67.

Two phase-advance units, respectively 73 and 74, receive the h signalfrom the barometric height sensor and the signal V from the airspeedsensor 17 and deliver respective signals It and V'. The action of theunit 73 is substantially equivalent to `differentiating and smoothingthe signal h, which is akin to a rate signal. To the extent that lz isgreater than k2, VTOC is made greater than V2, but with an upper limitat V3. The object is to ensure that the take-off path never falls belowthe scheduled take-oh. net climb path.

Further adjustments are provided respectively by units 75, 84 and 77 atkq, ka and kv. These are adjustments of overall gain in various parts ofthe system. The adjustment of kq sets the sensitivity of the system topitch rate changes and enables the amplitude of the pitch rate signalsto be adjusted to match the indicator scale. kv sets the sensitivity tochanges :in airspeed, specifically V', while ko sets the sensitivity tochanges in pitch angle.

The function unit 78 receives a bank angle signal rp from the attitudereference system 62 and also receives the signal V2 from the unit 65.The signal V2 is the takeoff safety speed below which the aircraft isnot permitted to climb, being variable with all-up weight etc., and isdetermined for specic conditions by the system.

The attitude reference system 62 provides the pitch attitude signal t9,the heading angle signal t and the bank angle signal 4i already referredto. The unit 79 provides a demanded heading angle signal il/D which iscombined with the heading angle signal 1]/ in a diiferencing unit 80 toprovide an output :,lf-i//D which is applied to a limiter 81. Thelimiter 81 is controlled by a heading error limit signal k5 provided bya heading error limit unit 82. The purpose of limiting the signal1,/J-\Al/D is to limit the bank angle demand, however great is therequired change of heading. The maximum bank angle is laid down for allaircraft.

The constant k3, provided by a unit 76, will be referred Ito in detaillater.

A further unit 83 provides an adjustment which governs the lateralsensitivity of the indicator index.

As in the system of FIGURE 2, a voltage source 33 is provided, togetherwith the port and starboard oleo leg switches 24 and 24a, connected inseries and closed when the respective legs become fully extended.

The constant AHR is a signal representing an amount of nose-up rotationsomewhat greater than the minimum value which will produce lift-olf.

A time delay unit 85 provides a time delay signal T, the purpose ofwhich is the same as the delay T1 of FIGURE 2.

The constant k3 from the unit 76 is fed to a phase advanced airspeederror computer 86, together with the signal from the function unit 7S,representing ,g/ V sin 15, tan tp, the V2 and V3 signals from the unit65, and the constant k1 from the unit 70. The constant k2 from the unit72 is combined with the signal lz' from the phase advance unit 73 in adiiferencing unit 87 and the resulting signal h-k2 is also fed to theunit 86, together with the signal V from the unit 74 and the constantkv. The output from the unit S6, which is in the form together with theconstant kc1 from the unit 75, is fed to a pitch demand computer 88,which also receives the rate or rate -ff rate signal from the gyroscope16 over a yline 89 together with the constants ko and A0E and the timedelay signal T from the unit 85. Operation of the computer 88 in phase Cis initiated by the signal from the switches 24 and 24a when both oleolegs become fully extended, subject to the time delay T from the startof phase B as in the case of the FIGURE 2 embodiment. The output of thecomputer S8 is fed to a director index pitch servo 90 and the output ofthe latter unit is used to actuate the director indicator index 91,shown in FIGURE 4. The signal from the gyroscope 16 is changed from therate to the rate i-lf rate form by a control signal from the pitchdemand computer 8S applied to the gyroscope over a line marked controlline.

The pitch attitude signal 0 is applied to a pitch attitude display servounit 92 which in turn feeds the indicator instrument of FIGURE 4.

The signal (il/-r//D) k5 from the limiter 81 is applied `to a rolldemand computer 93, which also receives the bank angle signal gb fromthe unit 62 and the constant signal 7W, 7% from the units 83 and 96 andfeeds a director index-roll servo 94, the output of which is alsoapplied to the indicator instrument of FIGURE 4.

The bank angle signal p is also applied to a bank attitude display servo95 the output of which is also applied to the indicator instrument ofFIGURE 4.

The sequence and method of operation of the system are described below,and in order to promote a ready understanding of the description, theswitching modes of the pitch demand computer and the form of outputsignal to the indicator instrument for the different phases are Agivenin tabular form.

l Phase C.-As described in relation to FIGURE 2, the aircraft incidenceis now further increased to -produce upward acceleration.

The phase is initiated by full extension of both port and starboard mainoleo legs closing microswitches, initia- TABLE Phase of Take-ofiConditions for Phase Engagement Ptl d(lyre Output of Pitch DemandComputer l o e N one Engaged Power on Rate -kqq (when gyro rotor has runup). A (Pre-Rotation) Power on, Director on, V VR Rate kv i(V-Vn)kqq. B(Rotation) Power on, Director on, V Vn Rate-lf Rate. ka (A0E-fd d t)kqq.C (Pull-up) and D (Initial Power on, Director on, Both Main Rate kv(V1-VToo) ks-kqq.

Climb). U/C Oleos Fully Extended and at least Time T after Phase BEngaged.

Power on.-When power is switched on the pitch demand computer passes thegyro rate signal to the director and the index immediately centres underthe action of the electrical spring on the gyro gimbal/rotor system.When the rotor has run up, the index will move about the zero demandmarker with aircraft pitch rate, the upward displacement being -kqq.

Director OIL- Before or during taxying to the take-olf starting pointthe director system is switched on, as in the case of the embodiment ofFIGURE 2, whereupon the pitch demand computer passes the kVf(V-VR)signal through to the director, making the total signal to the directornow kVf(V-VR)-kqq, where the function f(V-VR) is equal to V-VR exceptfor V VRW when the Value is limited to VRW- VR, the limit being imposedby the constant h4. The index is displaced downwardly by an amountcorresponding to the preset value of VR- VRW, to the marker S8.

Phase L -This phase, land the events and conditions associatedtherewith, are identical with those described in relation to theembodiment of FIGURE 2, and no further description is deemed necessary.

As in Director on the signal to the director is kvf V-VR -kqq PhaseB.--A rapid and chosen increase of incidence is now `required.

As the airspeed reaches VRlthe change of sign of the f(V- VR) signalinto the pitch demand computer simultaneously switches out this signaland switches the gyroscope 16 from the rate to the rate -i-f rate modeso that the index, now displaced from centre by an amount proportionalto k(A6R-fq)kqq, icks up to demand nose-up rotation of amount 0R at aninitial rate This pitch rate demand then falls 01T and by the time theindex is steady at centre with the mean pitch rate zero the aircraftwill have rotated up stably through the angle AGR. The value of AHR is,as previously stated, chosen somewhat greater than the value which willgive lift-off, with the object of obtaining lift-ott with a pitch rateSlightly less than that required in the early part of the pull-up, withthe proviso that the pitch rate at lift-off must not be so great as toresult in the aircraft tail-skid touching the runway.

The maximum director index displacement may be made to correspond to:about of pitch angle demand. With k=3ko this maximum displacement willalso correspond to 5/second pitch rate demand. So, taking a typical A0Bof 12, the index 91 will initially move fourfifths of the distance tothe top of the Scale at the start of Phase B onto the marker 59 shown inFIGURE 4 and this may also be considered as a demand for a 4/ secondpitch rate. During this phase the aircraft continues to accelerate landif MR has been correctly chosen the aircraft will leave the runway withan optimal nose-up pitch rate of about l to 11/2 /second, which isslightly less than that required in the iirst part of the pull-up.

tion being however prohibited until -at least time T after the start ofphase B; T will be of order 2 seconds. The pitch demand computerswitches the pitch gyroscope back from the rate -l-f rate to the ratemode and passes the signal UCV( V-VTOC) k3 through to the indicator.

The pull-up is thus demanded by displacing the indicator index upwardsby an amount proportional to [kV(V-VTOC)] c3-kqq. The initial pitch ratedemand is normally [c3/hq and is limited to some value of order 11/2/second which, in a steady .pull-up (incidence constant) at a true`airspeed of knots, gives about 1/6 g. The limiting value is determinedby the constant k3. When the phase-advanced speed demand signal later`comes ot its limit k3 the pilot is direct-ed to take off the pitch.rate and this should be accomplished in time governed by the timeconstant of phase-advance of V. Since this pitch rate of 11/2 second isonly slightly greater than the designed-for value at lift-oft the indexshould initially move slightly above centre, by an `amount proportionalto k3-kqq. The index should be central, or nearly so, at the end ofphase B, hence it will be expected to make an insignificant movement(probably upwards) on engagement of phase C. Thus, although the controllaws in phase B and phase C diier considerably this should not beapparent to the pilot, who will rotate to lift-oit and pull-up into theinitial climb as `a single manoeuvre,

Phase D Although no switching action occurs in the computer afterinitiation of phase C, this initial climb phase begins when the airspeedis within some small margin (say 2 or 3 knots) of VTOC (by which timethe undercarriage should be substantially retracted) and continues tothe start of the en-route climb at, perhaps, 1,500 feet height.

The normal objective is to hold the airspeed at the value V3, whichtypically exceeds V2 by :at least l0 knots, and possibly 30 or 40 knotsor more but if, abnormally, the value of h is below a certain value k2then the demanded speed is reduced by an amount proportional by thepreset constant k1 to this decit but with a lower limit of V2 when h thepreset constant value k2. In special circumstances, such as lack ofpower or deficient performance, the available power lshould be used forclimbing rather than accelerating. Although h' may differ somewhat fromrate of climb h it may be said that for nearly constant climb rate thetwo quantities will diiTer little and that the value k2 is closelyrelated to the climb rate appropriate to airspeed V2 in the scheduledtake-01T net climb path.

The value of the take-oilc climb speed defined in this way has beendenoted by VTOC and varies between the limits V2 and V3:

as in phase C.

The speeds V2 and V3 are in the nature of variable constants to theextent that they are generated as functions of both aircraft weight andflap angle, the Iformer being set in by the pilot and the latter sensedautomatically, as explained previously. The flap ang-le is sometimesreduced quite soon after take-off (but not below 400 feet under presentcivil aircraft regulations) and the value of VTOC will then usuallyincrease as the flaps are retracting, due to the consequent increases inV2 and V3; these increases will, however, be small, of the order 5knots. The director index will move appropriately to demand VTOC atevery instant, allowing also for the effects of iiap angle change on Vand h and of any trim change at constant stick position and airspeed.

The system operation in take-off and in initial climb with wings levelhas been described without consideration of either thrust loss orapplication of noise abatement procedures. Nor has the way in which thesystem directs the pilot to carry out baulked landing procedure beenconsidered. The operation of the system under these conditions will nowbe described. Lateral control of the aircraft will not be specificallydescribed since this is already well known.

Eect of thrust loss The system is designed to give optimal ornear-optimal short and long term performance in this important, thoughrare, eventuality.

Thrust loss during rotation or just after lift-off will have noimmediate effect on the indicator though the pull-up will be of shorterduration because the term kV(V-VTOC) will come off its limit k3 soonerdue to the reduction of V in spite of the possible reduction in VTOC.

Later thrust loss may result in a more or less rapid downward movementof the director index and the answering pilots forward stick movementwill check the reduction in airspeed which might otherwise occur.

The steady climb will be at a flatter angle than normal and usually at alower airspeed, possibly as low as V2, but exceeding V2 if the climbrate is sufiiciently high despite the subnormal thrust.

Noise abatement Baal/ced landing If on the landing approach theestimated aircraft weight is set up in the director system then, withpower on and director on and undercarriage down the system will beoperating in phase C (pull-up) since the airspeed will surely exceed theVR appropriate to this weight. The

vspeed demand VTOC will, because of the negative h', be

equal to V2.

The pilot initiates overshoot procedure by opening all throttles, fullyretracting the undercarriage and calling up take-off liap (notnecessarily in this order). It is also necessary to disengage theautopilot if in operation. The indicator index will certainly thendemand the limited pitch rate (typically 11/z/second) used in thepull-up at take-off because the aircraft is accelerating from anairspeed higher than the demanded airspeed.

Soon after establishing a climb rate the demanded airspeed VTOC mayincrease from V2 towards V3 which may reduce or even eliminate the speeddrop otherwise necessary to get to VTOC.

Extra pitch rate demand in turns Neglecting the effect of climb angle,which is small for the ltype of aircraft likely to use the system, theextra pitch rate demand in radians/second is g/ VT sin e tan qi, WhereVT is the true .airspeed. This is shown in FIGURE 4 approximated by g/V2 sin p tan rp Where V2 is C.A.S.

(corrected airspeed) or possibly I.A.S. (indicated airspeed).

Due to this the pitch signal to the director, in phases C and D, i.e.,when airborne, is

Two embodiments of the invention have been covered in `the foregoingdescription, and it will be clear that modifications of theseembodiments, or other embodiments, may be devised within the scope ofthe invention as defined in the appended claims.

What I claim is:

1. A take-off director system for aircraft comprising means forcontinuously measuring selected components of the movement and positionof the iaricraft, such means comprising a pitch rate gyroscope having arate output and an integrated rate output, and an airspeed sensor, meansto generate signals representing constants dependent upon thecharacteristics of the aircraft, means to combine the measurementssignals with the constants signals to provide a composite demand signalrepresentative of the action required to perform the 'take-off andinitial climb cycle in accordance with a predetermined program, and adirector indicating instrument to display the composite signal so thatthe pilot, by setting the controls to carry out the demands displayed bythe system, can perform the required sequence of take-off controloperations in a substantially optimal manner.

2. A system as claimed in claim 1 including further measuring -meanscomprising a barometric height sensor to measure the height of theaircraft, and an attitude reference system to measure the azimuth orheading angle, the pitch angle and the bank angle of the aircraft.

.3. A system as defined in claim 1 wherein the constants arek1representing a nose-up pitch rate which should not be exceeded duringthe early part of the take-off,

VR representing the predetermined airspeed lat which rotation of theaircraft may commence,

and MR representing a nose-up change in the aircrafts attitude slightlygreater than that which will produce lift-off.

4. A system as claimed in claim 3 including a further constant VRWrepresenting an airspeed below VR which is used in the indicatinginstrument to warn the pilot that the airspeed is approaching VR.

5. A system as claimed in claim 4 including further constants,

k3 representing an initial pitch-rate demand k representing (VR- VRW).

6. A system as `defined in claim 1 comprising means to set in a constantrepresenting the weight of the aircraft at the commencement of a flight.

7. A system as defined in claim 1 wherein the indicator instrumentcomprises an index which is moved upwardly from la Zero position toindicate a demand for nose-up pitch rate of the aircraft and downwardlyto indicate a demand for nose-down pitch rate.

8. A system as claimed in claim 7 comprising means to cause lateralmovement of the index to demand correction of spiral instability.

9. A system as defined in claim 1 wherein the predetermined programmecomprises a first phase, being the initial take-off run of the aircraftup to the rotation speed, the signal |applied to the indicatorinstrument including a function HV-VR), the phase ending when theairspeed is `equal to the rotation speed.

10. A system as claimed in claim 9 comprising automatic switching meanswhich change the signal applied to the indicator when the airspeedbecomes equal to the rotation speed.

11. A system as claimed in claim 10 in which a second phase comprisesthe continued run of fthe aircraft up to take-off, the signal applied tothe indicator showing a demand for upward rotation of the aircraft untiltake-off occurs.

12. A system as claimed in claim 11 comprising switch means associatedwith the oleo legs of the undercarriage tto change the signal to athird-phase signal demanding further upward rotation and increase ofspeed lup to the initial clim'b speed, comprising also time delay meansto delay the application of the third-phase signal to the indicator toprevent premature initiation of the third phase resulting from temporaryfull extension of the oleo legs (for example, due to unevenness in therunway) until the aircraft has been lifted clear of the ground.

13. A system as claimed in claim 12 in which the time delay means isarranged to prevent initiation of the third phase until a predeterminedtime from the commencement of the second phase.

14. A system as defined in clairn 1 4comprising a noiseabate switch foruse Iwhen the aircraft noise must be kept below a certain level byreducing the engine speed, and means controlled by the switch foraltering the value of at least one constant, thereby altering thepredetermined take-off program-Ine.

15. Apparatus for directing the take-off of an aircraft in successivephases, said apparatus comprising, in combination, means for generatinga first signal proportional to the rate of change of aircraft pitchangle and a second signal proportioned to the integral of such rate ofchange of aircraft pitch angle, means for generating a third signalIproportional to the difference between the instantaneous velocity ofthe` aircraft and a predetermined velocity at which the aircraft mustmove during the second phase, means operative during the first phase forsuperimposing said rst vand third signals to produce a first commandsignal, means for generating a fourth signal proportional to the`difference between a signal representing a predetermined pitch angle ofthe aircraft and said second signal, means operative during the secondphase for superimposing said first and fourth signals to produce asecond command signal, means for generating a fifth signal proportionalto the difference between the instantan-ecus velocity of the aircraftand a predetermined velocity at which the aircraft must move during thethird phase, means operative during the third phase for superimposingsaid first and fifth signals to produce a. third command signal, and adirector indicating instrument for visually displaying said commandsignals.

16. The apparatus according to claim 15 including means operative duringthe fourth phase for producing a command signal which is said fifthsignal.

References Cited by the Examiner UNITED STATES PATENTS 3,077,110 2/1963Gold 73--178 3,105,660 10/1963 Lenefsky et al. 244-77 3,148,540 9/1964Gold 73-178 3,174,710 3/1965 Hoekstra 73-178 X 3,200,642 8/ 1965Neuendorf et al 73-178 LOUIS R. PRINCE, Primary Examiner.

DONALD O. WOODIEL, Assistant Examiner.

1. A TAKE-OFF DIRECTOR SYSTEM FOR AIRCRAFT COMPRISING MEANS FORCONTINUOUSLY MEASURING SELECTED COMPONENTS OF THE MOVEMENT AND POSITIONOF THE AIRCRAFT, SUCH MEANS COMPRISING A PITCH RATE GYROSCOPE HAVING ARATE OUTPUT AND AN INTEGRATED RATE OUTPUT, AND AN AIRSPEED SENSOR, MEANSTO GENERATE SIGNALS REPRESENTING CONSTANTS DEPENDENT UPON THECHARACTERISTICS OF THE AIRCRAFT, MEANS TO COMBINE THE MEASUREMENTSSIGNALS WITH THE CONSTANTS SIGNALS TO PROVIDE A COMPOSITE DEMAND SIGNALREPRESENTATIVE OF THE ACTION REQUIRED TO PERFORM THE TAKE-OFF ANDINITIAL CLIMB CYCLE IN ACCORDANCE WITH A PREDETERMINED PROGRAM, AND ADIRECTOR INDICATING INSTRUMENT TO DISPLAY THE COMPOSITE SIGNAL SO THATTHE PILOT, BY SETTING THE CONTROLS TO CARRY OUT THE DEMANDS DISPLAYED BYTHE SYSTEM, CAN PERFORM THE REQUIRED SEQUENCE OF TAKE-OFF CONTROLOPERATIONS IN A SUBSTANTIALLY OPTIMAL MANNER.